1. Field of the Invention
The present invention relates to systems and methods for statically and dynamically balancing a mechanical structure. Particularly, the present invention relates to such systems and methods for space applications.
2. Description of the Related Art
Mass trim mechanisms are used to remove imbalance in a mechanical structure such as a spinning spacecraft, sensor, or other payload. Such imbalances can create disturbance torques that cause pointing errors in the highly sensitive antennas and other instruments on a spacecraft. Typically the spacecraft or payload is nominally balanced on the ground using balance weights. Imbalances evident on orbit and not removed during ground processing are referred to as “residual imbalances”. These residual imbalances can result from various sources including outgassing, 1-G sag (the effect of going from a 1-G build environment to a zero-G operational environment), balancing process uncertainty and deployment repeatability errors.
Consequently, residual static and dynamic imbalances are typically reduced or eliminated during on-orbit testing using mass trim mechanisms. Mass trim mechanisms can be used to adjust mass properties in order to correct both static imbalances, adjusting the center of gravity (CG), and dynamic imbalances, adjusting the products of inertia (POI). Balance corrections can also be made throughout the spacecraft life. For example, such adjustments can be necessary as mass properties of the spacecraft change as fuel is expended.
Thus, mass trim mechanisms are employed in spacecraft whenever it is necessary to adjust the mass properties of the vehicle. The center of gravity position is important to maintaining control of the satellite on orbit. For example, a single mass trim mechanism can be used on orbiting satellites to adjust the CG position of the spacecraft or a significant payload substructure, e.g. a sensor or other payload. In addition, when used in combination with one or more other mass trim mechanisms, a mass trim mechanism can be used to statically and dynamically balance an entire spacecraft, e.g. a spin stabilized satellite.
FIGS. 1A and 1B illustrate one example of a typical prior art mass trim mechanism that uses a leadscrew. As shown in FIG. 1A, the movable mass 100 is a single lumped element that can be translated in either direction along a slider rail 102 which supports the mass 100. The movable mass 100 is engaged with a leadscrew 104. Movement of the mass 100 is facilitated by driving the leadscrew 104 with a motor assembly 106 positioned at one end of the leadscrew 104. FIG. 1B illustrates the complexity of movable mass 100 assembly in requiring a large number of parts.
FIG. 1C illustrates another prior art mass trim mechanism. In this example, the movable mass is enclosed within the housing 108. The end of the leadscrew 104 is visible at end of the housing 108 and the motor assembly 106 to drive the leadscrew 104 is shown at the other end of the housing 108.
FIG. 1D illustrates yet another example of a prior art mass trim mechanism. In this case, the motor assembly 106 is attached to the movable mass 100 and effectively becomes part of the moveable mass 100. The motor assembly 106 and movable mass 100 together translate along the leadscrew 104 to adjust the spacecraft mass properties. Merging the motor assembly 106 with the movable mass effectively improves the overall mass efficiency of the design when compared to the designs of FIGS. 1A–1C, above. However, power and control of the mechanism must be brought to the movable mass 100 through wiring 110. This complicates the design of FIG. 1D and potentially undercuts the overall reliability.
The critical performance measurement of a mass trim mechanism can be defined as the amount of moveable mass times the maximum travel length. This figure indicates the maximum amount of influence that the mass trim mechanism can have on the mass properties of the attached mechanical structure. The additional components of the mechanism (i.e., everything except the movable mass) can be described as the mass overhead. All of the foregoing mass trim mechanisms use a mass that translates along a leadscrew 104. Such designs are inefficient in terms of the amount of movable mass and the travel distance relative to their total mass. Mass efficiency can be roughly defined as the movable mass divided by the total mass.
Another problem associated with conventional mass trim mechanisms is that they present problems being secured during launch. Prior to launch, the mass is typically positioned so that the center of gravity of the associated mechanical structure is nominally positioned. Accordingly, the movable mass is usually centered on the leadscrew, allowing the maximum adjustment in either direction after the satellite has achieved its orbit. In addition, if the mechanism fails during launch, negative impact on the mission is minimized. However, positioning the movable mass in the center of the leadscrew for launch presents some difficulty in securing the mechanism. The leadscrew and supporting structure must be sized to survive the launch loads while in this configuration. If longer travel or a larger mass are required, a larger and heavier leadscrew and support structure are needed. Alternately, an additional locking mechanism can be used to secure and support the movable mass during launch. In any case, these solutions add further to the overhead mass, complexity and cost of developing and delivering conventional mass trim mechanisms and potentially negatively impact reliability.
Accordingly, there is a need for mass trim mechanisms and methods which minimize mass overhead and increase mass efficiency. There is further a need for such mechanisms and methods which provide greater performance, i.e. more movable mass and/or a increased displacement distance. There is also a need for such mechanisms and methods which are simple, less costly and easily locked for launch. As detailed hereafter, these and other needs are met by the present invention.